Trailing edge cooling for a turbine blade

ABSTRACT

A component for a gas turbine engine comprises an airfoil having an outer surface. One or more cooling passages can be disposed within the airfoil, having a cooling passage extending along a trailing edge. A plurality of cooling channels can extend from the cooling passage through the trailing edge. At least one flow element and at least one film hole can be disposed in the cooling channel or the trailing edge passage adjacent the cooling channel. The flow element and the film hole can be in a predetermined relationship with one another providing improved flow to the film hole.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a Divisional Application of and claims priority toU.S. patent application Ser. No. 14/958,082, filed Dec. 3, 2015, nowU.S. Pat. No. 10,344,598, issued Jul. 9, 2019, the entirety of which isincorporated herein by reference. This application is also related toChinese Application 201611095392X, which was filed Dec. 2, 2016.

TECHNICAL FIELD OF THE INVENTION

This disclosure generally relates to a component or airfoil for aturbine engine which includes a predetermined relationship among a filmhole upstream and a trailing edge cooling channel, with the film holeprovided in a cooling passage upstream of the trailing edge coolingchannel.

BACKGROUND

Turbine engines, and particularly gas or combustion turbine engines, arerotary engines that extract energy from a flow of gases passing throughthe engine in a series of compressor stages, which include pairs ofrotating blades and stationary vanes, through a combustor, and then ontoa multitude of turbine blades. Gas turbine engines have been used forland and nautical locomotion and power generation, but are most commonlyused for aeronautical applications such as for airplanes, includinghelicopters. In airplanes, gas turbine engines are used for propulsionof the aircraft.

Gas turbine engines for aircraft are designed to operate at hightemperatures to maximize engine thrust, so cooling of certain enginecomponents, such as the rotor post is necessary during operation.Typically, cooling is accomplished by ducting cooler air from the highand/or low pressure compressors to the engine components, which requirecooling.

Flow elements placed on a surface complementary to a plurality of filmholes within the cooling flow can be utilized as a thermal coolingfeature, however, the flow elements can generate an unsteady flow as thecooling flow passes over them. The unsteady flow can provide an unsteadystream of fluid to the film holes reducing film cooling efficiency.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, the disclosure relates to a component for a gas turbineengine comprising an airfoil having an outer surface extendingchord-wise from a leading edge to a trailing edge and span-wise from aroot to a tip, a cooling passage located within the airfoil andextending along the trailing edge, at least one trailing edge coolingchannel extending from the cooling passage through the trailing edge,and at least one film hole having an inlet in the cooling passage, anoutlet on the outer surface, and a passage connecting the inlet and theoutlet. The inlet is located in the cooling passage in a predeterminedrelationship to the trailing edge cooling channel.

In another aspect, the disclosure relates to a turbine enginecomprising: a core engine comprising a casing at least partiallysurrounding a high pressure compressor, a combustor, and a high pressureturbine in a serial flow arrangement; an airfoil, provided in one of thecompressor or the combustor, having an outer surface extendingchord-wise from a leading edge to a trailing edge and span-wise from aroot to a tip; a cooling passage located within the airfoil andextending along the trailing edge; at least one trailing edge coolingchannel extending from the cooling passage through the trailing edge;and at least one film hole having an inlet in cooling passage, an outleton the outer surface, and a passage connecting the inlet and the outlet;wherein the inlet is located in the cooling passage in a predeterminedrelationship to the trailing edge cooling channel.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic, sectional view of a gas turbine engine.

FIG. 2 is a perspective view of an engine component in the form of aturbine blade of the engine of FIG. 1.

FIG. 3 is a cross-sectional view of the blade of FIG. 2 illustrating atrailing edge cooling channel.

FIG. 4 is a radial cross-section of the blade of FIG. 2 illustratingmultiple trailing edge cooling channels.

FIG. 5 is an enlarged view of the trailing edge of the blade of FIG. 3comprising flow elements within the cooling channel.

FIG. 6 is an enlarged view of the trailing edge illustrating flowelements on both sides of the cooling channel.

FIG. 7 is a further enlarged view of the trailing edge illustrating flowelements within a cooling passage.

FIG. 8 is a schematic side view of the trailing edge having a pluralityof film holes aligned with the cooling channels.

FIG. 9 is a schematic side view of the trailing edge having a pluralityof film holes offset from the cooling channels.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

The described embodiments of the present invention are directed toapparatuses, methods, and other devices related to routing airflow in aturbine engine. For purposes of illustration, the present invention willbe described with respect to an aircraft gas turbine engine. It will beunderstood, however, that the invention is not so limited and can havegeneral applicability in non-aircraft applications, such as other mobileapplications and non-mobile industrial, commercial, and residentialapplications.

It should be further understood that for purposes of illustration, thepresent invention will be described with respect to an airfoil for aturbine blade of the turbine engine. It will be understood, however,that the invention is not limited to the turbine blade, and can compriseany airfoil structure, such as a compressor blade, a turbine orcompressor vane, a fan blade, a strut, a shroud assembly, or a combustorliner or any other engine component requiring cooling in non-limitingexamples.

As used herein, the term “forward” or “upstream” refers to moving in adirection toward the engine inlet, or a component being relativelycloser to the engine inlet as compared to another component. The term“aft” or “downstream” used in conjunction with “forward” or “upstream”refers to a direction toward the rear or outlet of the engine relativeto the engine centerline.

Additionally, as used herein, the terms “radial” or “radially” refer toa dimension extending between a center longitudinal axis of the engineand an outer engine circumference.

Furthermore, as used herein, the terms “stream-wise” or “streamline,” orsimilar nomenclature when used with flow, fluid, gas, location, oralignment refers to a fluid or gas flow direction which can be linear ora vector of the flow where the flow is non-linear, where the directionof the flow is moving at any position or point in time.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise,upstream, downstream, aft, etc.) are only used for identificationpurposes to aid the reader's understanding of the present invention, anddo not create limitations, particularly as to the position, orientation,or use of the invention. Connection references (e.g., attached, coupled,connected, and joined) are to be construed broadly and can includeintermediate members between a collection of elements and relativemovement between elements unless otherwise indicated. As such,connection references do not necessarily infer that two elements aredirectly connected and in fixed relation to one another. The exemplarydrawings are for purposes of illustration only and the dimensions,positions, order and relative sizes reflected in the drawings attachedhereto can vary.

FIG. 1 is a schematic cross-sectional diagram of a gas turbine engine 10for an aircraft. The engine 10 has a generally longitudinally extendingaxis or centerline 12 extending forward 14 to aft 16. The engine 10includes, in downstream serial flow relationship, a fan section 18including a fan 20, a compressor section 22 including a booster or lowpressure (LP) compressor 24 and a high pressure (HP) compressor 26, acombustion section 28 including a combustor 30, a turbine section 32including a HP turbine 34, and a LP turbine 36, and an exhaust section38.

The fan section 18 includes a fan casing 40 surrounding the fan 20. Thefan 20 includes a plurality of fan blades 42 disposed radially about thecenterline 12. The HP compressor 26, the combustor 30, and the HPturbine 34 form a core 44 of the engine 10, which generates combustiongases. The core 44 is surrounded by core casing 46, which can be coupledwith the fan casing 40.

A HP shaft or spool 48 disposed coaxially about the centerline 12 of theengine 10 drivingly connects the HP turbine 34 to the HP compressor 26.A LP shaft or spool 50, which is disposed coaxially about the centerline12 of the engine 10 within the larger diameter annular HP spool 48,drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20.The portions of the engine 10 mounted to and rotating with either orboth of the spools 48, 50 are referred to individually or collectivelyas a rotor 51.

The LP compressor 24 and the HP compressor 26 respectively include aplurality of compressor stages 52, 54, in which a set of compressorblades 58 rotate relative to a corresponding set of static compressorvanes 60, 62 (also called a nozzle) to compress or pressurize the streamof fluid passing through the stage. In a single compressor stage 52, 54,multiple compressor blades 56, 58 can be provided in a ring and canextend radially outwardly relative to the centerline 12, from a bladeplatform to a blade tip, while the corresponding static compressor vanes60, 62 are positioned downstream of and adjacent to the rotating blades56, 58. It is noted that the number of blades, vanes, and compressorstages shown in FIG. 1 were selected for illustrative purposes only, andthat other numbers are possible. The blades 56, 58 for a stage of thecompressor can be mounted to a disk 53, which is mounted to thecorresponding one of the HP and LP spools 48, 50, with each stage havingits own disk. The vanes 60, 62 are mounted to the core casing 46 in acircumferential arrangement about the rotor 51.

The HP turbine 34 and the LP turbine 36 respectively include a pluralityof turbine stages 64, 66, in which a set of turbine blades 68, 70 arerotated relative to a corresponding set of static turbine vanes 72, 74(also called a nozzle) to extract energy from the stream of fluidpassing through the stage. In a single turbine stage 64, 66, multipleturbine blades 68, 70 can be provided in a ring and can extend radiallyoutwardly relative to the centerline 12, from a blade platform to ablade tip, while the corresponding static turbine vanes 72, 74 arepositioned upstream of and adjacent to the rotating blades 68, 70. It isnoted that the number of blades, vanes, and turbine stages shown in FIG.1 were selected for illustrative purposes only, and that other numbersare possible.

In operation, the rotating fan 20 supplies ambient air to the LPcompressor 24, which then supplies pressurized ambient air to the HPcompressor 26, which further pressurizes the ambient air. Thepressurized air from the HP compressor 26 is mixed with fuel in thecombustor 30 and ignited, thereby generating combustion gases. Some workis extracted from these gases by the HP turbine 34, which drives the HPcompressor 26. The combustion gases are discharged into the LP turbine36, which extracts additional work to drive the LP compressor 24, andthe exhaust gas is ultimately discharged from the engine 10 via theexhaust section 38. The driving of the LP turbine 36 drives the LP spool50 to rotate the fan 20 and the LP compressor 24.

Some of the ambient air supplied by the fan 20 can bypass the enginecore 44 and be used for cooling of portions, especially hot portions, ofthe engine 10, and/or used to cool or power other aspects of theaircraft. In the context of a turbine engine, the hot portions of theengine are normally downstream of the combustor 30, especially theturbine section 32, with the HP turbine 34 being the hottest portion asit is directly downstream of the combustion section 28. Other sources ofcooling fluid can be, but is not limited to, fluid discharged from theLP compressor 24 or the HP compressor 26.

FIG. 2 is a perspective view of an engine component in the form of oneof the turbine blades 68 of the engine 10 from FIG. 1. The turbine blade68 includes a dovetail 76 and an airfoil 78. The dovetail 76 can beconfigured to mount to a turbine rotor disk on the engine 10. Theairfoil 78 extends from a tip 80 to a root 82 defining a span-wisedirection. The dovetail 76 further includes a platform 84 integral withthe airfoil 78 at the root 82, which helps to radially contain theturbine airflow. The dovetail 76 comprises at least one inlet passage,exemplarily shown as a first inlet passage 88, a second inlet passage90, and a third inlet passage 92, each extending through the dovetail 76to provide internal fluid communication with the airfoil 78 at a passageoutlet 94. The inlet passages 88, 90, 92 as shown are exemplary shouldnot be understood as limiting. More or less inlet passages can be usedto provide a flow of fluid internal of the airfoil 78. It should beappreciated that the dovetail 76 is shown in cross-section, such thatthe inlet passages 88, 90, 92 are housed within the body of the dovetail76. It should be further appreciated that as described herein, theengine component is described as an airfoil 78, however, this should notbe construed as limiting and additional engine components such as ablade, vane, strut, or shroud assembly, in non-limiting examples, can besubstituted for the airfoil.

Turning to FIG. 3, the airfoil 78, shown in cross-section, has an outerwall defining a concave-shaped pressure wall 98 and a convex-shapedsuction wall 100 which are joined together to define an airfoil shape. Aleading edge 102 and a trailing edge 104 define a chord-wise directionextending therebetween. The airfoil 78 rotates in a direction such thatthe pressure wall 98 follows the suction wall 100. Thus, as shown inFIG. 3, the airfoil 78 would rotate upward toward the top of the page.

The airfoil 78 comprises an interior 96 defined by a first coolingpassage 110, a second cooling passage 112, and a trailing edge coolingpassage 114. The trailing edge cooling passage 114 comprises a trailingedge cooling channel 116 extending from the trailing edge coolingpassage 114 through the pressure sidewall 98 adjacent the trailing edge104. Alternatively, the cooling channel 116 can extend through thetrailing edge 104 or the suction sidewall 100. A flow of fluidstream-wise gas S, such as a cooling fluid, can pass from the trailingedge cooling passage 114 through the cooling channel 116 and exhaust atthe trailing edge 104 of the airfoil 78.

In FIG. 4, illustrating a radial cross-section of the airfoil 78, theairfoil comprises multiple cooling channels 116, defined within atrailing edge wall 117. One can appreciate that the trailing edgepassage 114 is in fluid communication with the third inlet passage 92 atthe passage outlet 94. The trailing edge passage 114 feeds the pluralityof cooling channels 116, and exhausts any remaining gas through anexhaust outlet 121 at a tip flag 115. Furthermore, an air fed toward thetip 80 can exhaust into a tip channel 119, which can join with the gasexhausting from the exhaust outlet 121 at the tip flag 115. It should beunderstood that the cooling configuration shown in FIG. 4 is exemplary,and should not be understood as limiting.

Turning to FIG. 5, the cooling channel 116 comprises a plurality of flowelements 120 and a plurality of film holes 122. The flow elements 120can comprise turbulators, pins or pin banks, or mesh in non-limitingexamples. The flow elements 120 can be discrete members, with multipleflow elements 120 disposed radially along the cooling channel 116, orcan be a single elongated member extending along a portion of or theentire radial length of the cooling channel 116. The film holes 122 canhave an inlet 124 disposed in the cooling channel 116, an outlet 126disposed on the pressure sidewall 98 and a passage 128 connecting theinlet 124 to the outlet 126. While the film holes 122 are illustrated onthe pressure side 98, they can alternatively be placed on the suctionside 100.

The flow elements 120 are disposed across from the film holes 122 suchthat the inlets 124 for the film holes 122 are located in the coolingchannel 116 in a predetermined relationship to the flow element 120. Thepredetermined relationship comprises the inlet 124 of the film hole 122being located on an opposite side of the cooling channel 116 from theflow elements 120. The flow elements 120 are disposed on the wall of thecooling channel 116 in the same stream-wise location. Alternatively, thepredetermined relationship can be defined by the same location basedupon the centerline of the cooling channel 116 rather than thestreamline flow S. The stream-wise location can be defined as thedistance along the cooling channel 116 in the direction of thestream-wise flow S through the cooling channel 116. Additionally, thefilm holes 122 and the flow elements 120 can be arranged in pairs, suchthat the predetermined relationship comprises a pair of one film hole122 and one flow element 120.

It should be understood that the number of film holes and flow elementsare exemplary. There can be more or less film holes and flow elementsthan as shown. Furthermore, there need not be the same number of filmholes and flow elements.

Turning now to FIG. 6, an additional example illustrates additional flowelements 125 disposed on the wall of the cooling channel 116. Theadditional flow elements 125 are disposed between the film hole inlets124. It should be appreciated that flow elements 125 on the samesidewall as the film holes 122 need not be utilized in combination withthe flow elements 120 opposite of the film holes 122.

FIG. 7 illustrates another example having a plurality of film holes 130disposed in the trailing edge passage 114 upstream of the coolingchannel 116 relative to the stream-wise flow. There are two exemplaryfilm holes 130 shown, one on the pressure sidewall 98 and one on thesuction sidewall 100. It should be appreciated that the position andgeometry of the film holes 130 are exemplary. The film holes can beplaced nearer or farther from the cooling channel 116, and can comprisea plurality of film holes 130 disposed radially along the length of theairfoil 78. Each film hole 130 comprises an inlet 132 disposed in thetrailing edge passage 114 and an outlet 134 disposed on the outersurface, such as the pressure or suction sidewalls 98, 100. A passage136 fluidly couples the inlet 132 to the outlet 134.

A plurality of flow elements 140 are disposed within the trailing edgepassage 114. The flow elements 140 are disposed opposite of the filmholes 130 and can be disposed across from a film hole 130 in apredetermined relationship such that the flow elements 140 are spacedstream-wise from one another upstream of the cooling channel 116. Twostream-wise axes 138 are illustrated, disposed orthogonal to the streamwise flow S, such that the film hole 130 and the associated flow element140 are aligned in the predetermined relationship relative to thestream-wise flow S. Alternatively, the predetermined relationship can berelative to a centerline of the cooling channel 116. Additionally, itshould be appreciated that the flow elements as shown in FIG. 7 areoptional, and can comprise a trailing edge passage configuration withonly the film holes 130 or some of the flow elements 140.

It should be appreciated that the flow elements 140 can be multiple flowelements 140 disposed radially along the surface of the trailing edgepassage 114. Additionally, there can be multiple complementary filmholes associated with the multiple flow elements.

Turning now to FIG. 8, a radial schematic of the trailing edge passage114 and the trailing edge cooling channels 116 best illustrates multipletrailing edge cooling channels 116 disposed radially along the trailingedge of the airfoil 78. The plurality of film holes 130 can be disposedradially along the trailing edge passage 114 aligned with the trailingedge cooling channels 116. As such, an airflow entering the trailingedge cooling channels 116 will feed the film hole 130 prior to enteringthe trailing edge cooling channels 116.

Alternatively, in FIG. 9, the film holes 130 can be disposed offset fromthe trailing edge cooling channel 116. Thus, a flow of air entering thetrailing edge cooling channels 116 will not be disrupted by the airflowsentering the film holes 130.

Thus, the placement of the film holes 130 as shown in FIGS. 8 and 9 canbe utilized in combination with the alignments of the flow elements andthe film holes 130 as shown in FIG. 7 to place film holes 130 near thetrailing edge relative to both a cooling channel 116 and the flowelement 140.

It should be appreciated that the relative geometric placement of thefilm hole inlets to the airfoil trailing edge features, such as the flowelements, which can comprise turbulators, pins or pin banks, or mesh,can be beneficial to the film hole inlet flows and the film holedischarge coefficients. Additionally the placement of the flow elementsin the channel entry or upstream of the channel can facilitate flowentry into the film holes or the cooling channels. The film holes can besubstantially lined up with the placement of the flow elements togenerate the beneficial flow.

This written description uses examples to disclose the invention,including the best mode, and to enable any person skilled in the art topractice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they have structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A component for a turbine engine comprising: anairfoil having an outer surface extending chord-wise from a leading edgeto a trailing edge and span-wise from a root to a tip; a cooling passagelocated within the airfoil and extending along the trailing edge; atleast one trailing edge cooling channel extending from the coolingpassage to the outer surface at the trailing edge, defining astream-wise flow through the at least one trailing edge cooling channel;and at least one film hole having an inlet in the cooling passage, anoutlet on the outer surface, and a passage connecting the inlet and theoutlet; and at least one turbulator provided in the cooling passage;wherein the at least one turbulator and the inlet of the at least onefilm hole are spaced in the cooling passage in a predeterminedrelationship defining an axis extending between the at least oneturbulator and the inlet of the at least one film hole, with the axisarranged orthogonal to the stream-wise flow.
 2. The component of claim 1wherein the inlet is aligned with the at least one trailing edge coolingchannel relative to the span-wise direction.
 3. The component of claim 2wherein the inlet of the at least one film hole is aligned with acenterline of the at least one trailing edge cooling channel relative tothe stream-wise flow.
 4. The component of claim 1 wherein the inlet isstream-wise non-aligned with the at least one trailing edge coolingchannel relative to the stream-wise flow.
 5. The component of claim 1wherein the component is one of a rotating blade or a stationary vane.6. The component of claim 1 wherein the at least one trailing edgecooling channel comprises multiple trailing edge cooling channels. 7.The component of claim 6 wherein the inlet of the at least one film holeis located between two adjacent trailing edge cooling channels of themultiple trailing edge cooling channels in the span-wise direction. 8.The component of claim 7 wherein the at least one film hole comprisesmultiple film holes and the inlets of the multiple film holes arelocated between pairs of adjacent trailing edge cooling channels of themultiple trailing edge cooling channels.
 9. A turbine engine comprising:a core engine comprising a casing at least partially surrounding a highpressure compressor, a combustor, and a high pressure turbine in aserial flow arrangement; an airfoil, provided in one of the compressoror the combustor, having an outer surface extending chord-wise from aleading edge to a trailing edge and span-wise from a root to a tip; acooling passage located within the airfoil and extending along thetrailing edge; at least one trailing edge cooling channel extending fromthe cooling passage through the trailing edge to define a stream-wiseflow through the at least one trailing edge cooling channel; at leastone film hole having an inlet in the cooling passage, an outlet on theouter surface, and a passage connecting the inlet and the outlet; and atleast one turbulator provided in the cooling passage; wherein the inletof the at least one film hole is located in the cooling passage in apredetermined relationship to align with the at least one turbulator todefine an axis between the at least one turbulator and the inlet of theat least one film hole, with the axis being arranged orthogonal to thestream-wise flow.
 10. The turbine engine of claim 9 wherein the inlet ofthe at least one film hole is aligned with the at least one trailingedge cooling channel relative to the stream-wise flow.
 11. The turbineengine of claim 10 wherein the inlet of the at least one film hole isaligned with a centerline of the at least one trailing edge coolingchannel relative to the stream-wise flow.
 12. The turbine engine ofclaim 9 wherein the inlet is non-aligned with the at least one trailingedge cooling channel relative to the stream-wise flow.
 13. The turbineengine of claim 9 wherein the airfoil is one of a rotating blade or astationary vane.
 14. The turbine engine of claim 9 wherein the at leastone trailing edge cooling channel comprises multiple trailing edgecooling channels.
 15. The turbine engine of claim 14 wherein the inletis located between adjacent trailing edge cooling channels of themultiple trailing edge cooling channels.
 16. The turbine engine of claim15 wherein the at least one film hole comprises multiple film holes andthe inlets are located between adjacent trailing edge cooling channels.17. The turbine engine of claim 14 wherein the at least one film holecomprises multiple film holes and the inlets are located in stream-wisealignment with the trailing edge cooling channels.
 18. The turbineengine of claim 9 wherein the airfoil is a blade.